1. Field of the Invention
This invention relates to cooling of turbine airfoils and, more particularly, to bonded hollow turbine vanes having cooling inserts within outer airfoil walls.
2. Description of Related Art
It is well known to cool parts using heat transfer across walls having hot and cold surfaces by flowing a cooling fluid in contact with the cold surface to remove the heat transferred across from the hot surface. Among the various cooling techniques presently used are convection, impingement and film cooling as well as radiation. These cooling techniques have been used to cool gas turbine engine hot section components such as turbine vanes and blades. A great many high pressure turbine (HPT) vanes, and particularly the high pressure turbine inlet guide vane, also known as the combustor nozzle guide vane, utilize some form of a cooled hollow airfoil. An airfoil typically has a hollow body section which includes a leading edge having a leading edge wall followed by a pressure side wall and a suction side wall which form a substantial part of the outer wall which includes the hot wetted surface on the outside of the walls. The pressure and suction side walls typically converge to form a trailing edge.
Typically, a vane having a hollow airfoil is cooled using two main cavities, one with coolant air fed from an inboard radial location and the other with coolant air fed from an outboard location. These cavities typically contain impingement inserts which serve to receive cooling air and direct the coolant in impingement jet arrays against the outer wall of the airfoil's leading edge and pressure and suction side walls to transfer energy from the walls to the fluid, thereby, cooling the wall. These inserts are positioned by inward protrusions from the outer wall of the airfoil and are referred to as floating because they are not connected to or bonded to the outer wall. These protrusions or positioning dimples are integral with either the insert of the airfoil and provide the barest of contact between the insert and the airfoil wall and therefore are not very effective for forming convective cooling air passages between the insert and the airfoil wall. It is fairly well known in the prior art to use inserts for impingement cooling. However such designs are subject to stacking problems because inserts are inserted into the airfoil cavities after the airfoil has been formed, whether the airfoil is formed as a single piece casting or bonded together from two halves or sections of an airfoil that are cast or formed in some other manner. This limits the degree of twist and curvature variation of the airfoil in the radial direction.
Other prior art designs included double wall outer shell airfoils where the outer shell of the airfoil had an inner and outer wall integrally formed or bonded together to form cooling passages therebetween. Such designs are subject to large temperature differentials .DELTA.T across the airfoils shells thereby causing thermal stresses that could break the bond or separate the inner and outer walls. This, in turn, would reduce the structural integrity and effectiveness of the cooling passages and could lead to airfoil failure.
Another drawback of turbine airfoils using inserts disclosed in the prior is that the insert must be installed into the vane or airfoil cavity by inserting it through either a radially inner or outer diameter cavity opening. This imposes restrictions on the aerodynamics and airfoil stacking by requiring that the vane cavities be relatively straight (line of sight) with little if any twist or centerline curvature permissible. In addition to providing clearance to allow the insert to be inserted, additional clearance may be required to account for manufacturing tolerances on both the insert and cavity contour which further restrict the shapes of cooled airfoil designs.
Turbine vane cooling requires a great deal of cooling fluid flow which typically requires the use of power and is therefore generally looked upon as a fuel efficiency and power penalty in the gas turbine industry. Any improvement to the overall efficiency and effectiveness of turbine vane cooling would provide a great cost saving and fuel efficiency benefit to gas turbine designs. Therefore, there is a great need for a cooled airfoil design particularly for use in turbine vanes which have twist and/or a curved radially extending stacking line or centerline.
The present invention provides improved turbine airfoil cooling and engine efficiency and is particularly useful for cooled turbine vanes having airfoils which have twist and/or a curved radially extending stacking line or centerline.